PDF/ocr_dataset_1061.jsonl

{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 13, "total_pages": 46, "image_filename": "19930085548_p13.jpg", "text": "12\nNACA RM No. E8L30\n\ntime did not permit optimizing injection pattern, injection advance,\nor duration of injection. Furthermore, some loss in efficiency can\nbe attributed to the low charging efficiency. For these reasons,\nthe data of figure 19 are considered to be a reasonable check on the\nassumptions in the initial analysis, particularly at fuel-air ratios\nless than 0.03.\n\nExhaust Measurements\n\nHeat losses. - The heat loss to the coolant and the exhaust-\ngas temperature of the experimental cylinder are shown in figure 20\nas functions of fuel-air ratio. The heat losses decrease with an\nincrease in fuel-air ratio. This variation probably results because\nthe heat input increases faster than the temperature difference\nleading to heat transfer. All the data are considerably higher than\nthe value of 18 percent assumed in the previous analysis (refer-\nence 1). The higher heat rejection of the experimental cylinder is\npartly attributed to the high surface-volume ratio of this cylinder\n(the experimental cylinder has a surface-volume ratio over $2\\frac{1}{2}$ times\nthat of a 6.3- by 6.3-in. loop-scavenged two-stroke-cycle cylinder\n(Klöckner Humbolt Deutz engine, reference 8)), and partly to the\nlow coolant temperature used to expedite the investigation.\n\nExhaust-gas temperature. - The exhaust-gas temperature (fig. 20)\nis practically a linear function of fuel-air ratio. The slight amount\nof upward curvature is caused by a decrease in the engine efficiency\nas the fuel-air ratio is increased. The difficulty of attaining\nequilibrium conditions in the exhaust tank, which had considerable\nthermal lag, may account for the scatter in the data points. Because\nthe data were taken in the direction of increasingly rich mixtures,\nthe higher temperatures at each respective fuel-air ratio are con-\nsidered most valid and the line through the points is drawn accord-\ningly.\n\nComparison of calculated and experimental exhaust-gas temper-\natures. - The equation used in the previous analysis (reference 1)\nwas modified to eliminate the necessity for knowing the compressor-\ninlet temperature in the gas-generator engine, to eliminate the\nsimplifying assumption of constant specific heat, and to consider\nthe higher heat losses in the experimental cylinder. In its modified\nform, the equation becomes\n\n$$H_g = \\frac{(1 - q_l - \\eta_t) h_c \\left(\\frac{F}{A}\\right)}{\\left(1 + \\frac{F}{A}\\right)} + H_m \\quad (5)$$", "timestamp": "2026-07-22T06:17:06.709883+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 4, "total_pages": 60, "image_filename": "19930085862_p4.jpg", "text": "2\nNACA RM No. L9A07\n\nINTRODUCTION\n\nThe contemplated use of swept wings incorporating sharp-edged airfoil sections for high-speed airplanes has resulted in a need for information concerning the effectiveness of lateral-control devices on wings of this type. An investigation at low air speeds, therefore, has been made in the Langley 19-foot pressure tunnel to determine the lateral characteristics of a 42° sweptback wing which had sharp-edged, symmetrical, circular-arc airfoil sections and was equipped with either a conventional aileron or various spanwise arrangements of step spoilers.\n\nThe rolling-moment characteristics of the aileron and the spoilers together with the aileron hinge-moment, normal-force, and balance-chamber pressure characteristics were determined for both the plain wing and the wing equipped with various high-lift and stall-control devices. These devices included extensible, round-nose, leading-edge flaps, leading-edge drooped-nose flaps, trailing-edge split flaps, and upper-surface fences.\n\nThe investigation was conducted at Reynolds numbers ranging between $5.3 \\times 10^6$ and $6.9 \\times 10^6$ which corresponded to a Mach number range of 0.11 to 0.15.\n\nSYMBOLS\n\nThe data are referred to a set of axes coinciding with the wind axes and originating in the plane of symmetry at the quarter-chord point of the mean aerodynamic chord. All wing coefficients are based upon the dimensions of the basic wing.\n\n| | |\n| :--- | :--- |\n| $C_L$ | lift coefficient (Lift/qS) |\n| $C_{L_{max}}$ | maximum lift coefficient |\n| $C_D$ | drag coefficient (Drag/qS) |\n| $C_m$ | pitching-moment coefficient (Pitching moment/qSc) |\n| $C_n$ | yawing-moment coefficient (Yawing moment/qSb) |\n| $C_l$ | rolling-moment coefficient (Rolling moment/qSb) |", "timestamp": "2026-07-22T06:17:06.948771+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 13, "total_pages": 37, "image_filename": "19930082646_p13.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:17:08.021745+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 27, "total_pages": 65, "image_filename": "19930082546_p27.jpg", "text": "26\nNACA TN No. 1870\n\nquantity for a typical fuselage having $\\frac{c}{C_c} = 0.10$, M = 0.8 grams per centimeter$^2$, $\\omega_n = 2\\pi60 = 376$ radians per second, and $k = \\rho c = 42$ grams per centimeter$^2$-second is\n\n$$\n\\frac{c}{C_c} \\frac{M\\omega_n}{K} = \\frac{(0.10)(0.80)(376)}{42} = 0.70\n$$\n\nEquation 10 shows that the damping is effective in reducing resonant peaks for high values of $\\omega_n$ (high rigidity), mass, and damping coefficients. This explains why damping reduces the amplitude of the higher responses but is not very effective in reducing the low-frequency peaks.\n\nThe transmission coefficient $T_c$ of sound energy through a wall is given by the square of the ratio of wall amplitude to the amplitude of the impinging wave. The reciprocal of the transmission is given for the case of zero structural damping in reference 5 as\n\n$$\n\\frac{1}{T_c} = \\frac{\\left(M\\omega_1 - \\frac{s}{\\omega_1}\\right)^2}{4\\rho^2c^2} + 1\n$$\n\nwhere M is the mass of the wall per unit area, s is the stiffness ($s = M\\omega_n^2$ where $\\omega_n$ is natural frequency of panel), $\\omega_1$ is angular frequency of impinging sound, and c is velocity of sound in air.\n\nThis equation may be written for air at standard conditions (15° C and 760 mm of Hg) as\n\n$$\nT_c = \\frac{7056}{7056 + 4\\pi^2f_1^2M^2\\left(1 - \\frac{f_o^2}{f_1^2}\\right)^2} \\quad (11)\n$$\n\nwhere $f_1$ is the frequency of the impinging sound, $f_o$ the natural frequency of the fuselage, and M the mass per unit area of the fuselage.", "timestamp": "2026-07-22T06:17:08.882178+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 22, "total_pages": 28, "image_filename": "19930082703_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:17:09.697034+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 48, "total_pages": 114, "image_filename": "19930086061_p48.jpg", "text": "44\nNACA RM L9J07\n\nLeft semispan\nRight semispan\nUpper\nLower\nP\n(a) $\\psi = 0^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\nP\n(b) $\\psi = 10^\\circ$\n\nFigure 12.- Pressure distribution about wing 1 at various angles of yaw;\n$\\alpha = 4.1^\\circ$.", "timestamp": "2026-07-22T06:17:11.330875+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 31, "total_pages": 44, "image_filename": "19930082566_p31.jpg", "text": "NACA TN No. 1889\n29\n\n[Figure: Specimen in position for testing.]\n\nFigure 11.- Specimen in position for testing.", "timestamp": "2026-07-22T06:17:16.874769+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 15, "total_pages": 72, "image_filename": "19930085491_p15.jpg", "text": "14 CONFIDENTIAL NACA RM No. A5J04\n\nat approximately 33 percent of the normal-section chord. For this type of pressure distribution, reference 15 indicates that laminar-flow separation will occur when the kinetic energy is reduced to approximately 90 percent of its maximum value. This relationship may be written\n\n$\\frac{1-P}{1-P_{\\max }} \\cong 0.90$\n\nwhere P is the pressure coefficient at the point of laminar separation.\n\nAlthough the theoretical considerations of reference 16 excluded the effects of compressibility, the results of reference 18 indicate that the similarities between boundary-layer flow at low and transonic speeds justify the extension of the separation criterion to the present case where the leading-edge normal flow Mach number is approximately 0.7. Therefore, equation (5) has been applied directly to the normal-section pressure distributions of figure 6 to obtain the theoretical line of laminar separation.\n\nThe foregoing discussion which assumes that the subsonic separation criterion can be used in a supersonic flow field also neglects the fact that since the wing is tapered it is not an oblique cylinder as was used in reference 17. However, the pressure distributions are approximately two-dimensional inboard from the tip Mach cones and therefore application of the method is justifiable in this region. Near the wing root the pressure field is essentially three-dimensional and is considerably affected by wing-fuselage interference so that the theoretical line is of questionable accuracy in this region.\n\nEXPERIMENTAL RESULTS\n\nForce Tests\n\nThe results of the force tests are presented in the usual manner as lift, drag, and pitching-moment coefficients. The following tabulation summarizes the test conditions and figure numbers in which the results are presented:\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:17:19.455030+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 30, "total_pages": 50, "image_filename": "19930082592_p30.jpg", "text": "NACA TN 1914\n29\n\n[Figure: Micrograph showing a circular cross-section of a material. Labels point to the \"Oxidation interface\" at the top dark region and the \"Grain-boundary phase\" within the lighter matrix. A NACA logo with \"C-22908\" and \"2-7-49\" is in the bottom right corner of the image.]\n\nFigure 9. - Region adjacent to oxidation interface of 30-percent-molybdenum - titanium carbide ceramal. Grain boundary does not appear to contain appreciable quantity of grain-boundary phase. Temperature, $1785^\\circ$ F; time at temperature, 7 hours; etchant, potassium hydroxide plus potassium ferricyanide $\\text{KOH}+\\text{K}_3\\text{Fe(CN)}_6$; magnification, X1000.", "timestamp": "2026-07-22T06:17:20.230946+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 28, "total_pages": 46, "image_filename": "19930082613_p28.jpg", "text": "NACA TN 1938\n27\n\n[Figure: Micrograph showing a crack structure. Labels point to \"Main crack\", \"Tubular formation\", and \"Detail A\".]\n\n(a) Magnification, X60.\n\nNACA\nC-22651\n12-10-48\n\nFigure 8. - Portion of large crack showing surface and subsurface scale. Type-A liner; etchant, none; condition, failed in service.", "timestamp": "2026-07-22T06:17:21.648556+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 15, "total_pages": 17, "image_filename": "19930085572_p15.jpg", "text": "NACA RM No. E8L02\n13\n\nCorrected specific fuel consumption based\non net thrust, lb/(hr)/(lb thrust)\n\nPressure\naltitude\n(ft)\nO 5,000\n□ 10,000\n◇ 20,000\n△ 30,000\n\nFuel\n—O— AN-F-58\n--●-- AN-F-32\n\n[Figure: Graph plotting Corrected specific fuel consumption against Corrected engine speed]\n\nCorrected engine speed, rpm\n\nNACA\n\nFigure 4. - Comparison of corrected specific fuel consumption based\non jet thrust with AN-F-58 and AN-F-32 fuels. Mach number, 0.37.", "timestamp": "2026-07-22T06:17:22.511829+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 8, "total_pages": 31, "image_filename": "19930085859_p8.jpg", "text": "6\nNACA RM No. L9B25\n\nLift and Drag Characteristics\n\nThe isolated wing lift-curve slope measured near zero lift was about 0.066 at a Mach number of 0.60. (See fig. 7.) This compares with a value of 0.063 estimated for this Mach number by use of the charts in reference 3. In the Mach number range between 0.85 and 0.98 it appears that the maximum lift coefficient may be fairly close to 0.6 (fig. 7). The basic lift-curve slope was increased by an average of about 9 percent by the addition of the fuselage.\n\nThe drag rise at zero lift (fig. 13) began at a Mach number of about 0.89 for both the wing and wing-fuselage configurations. It is interesting to note that although this drag rise occurred at a Mach number about 0.04 lower than for the 45° sweptback wing (reference 1), which, except for sweepback, had geometric characteristics identical to those of the present wing, the values of $C_{D_{L=0}}$ and $(L/D)_{max}$ at the highest Mach numbers are not materially different for the two models. The absolute drag coefficients are probably high because of the presence of end-plate tares and the relatively low Reynolds numbers at which these tests were made.\n\nThe lateral center of pressure for the wing alone ($C_L = 0.4$) was located at 44 percent of the semispan at a Mach number of 0.6. This value compared with an estimated low-speed value of about 45 percent semispan (reference 3). Between M = 0.9 and 1.00 there was a fairly abrupt movement of $y_{c.p.}$ to about 50 percent semispan. This same outboard shift was obtained with the 45° sweptback wing at a somewhat higher Mach number. (See reference 1.) The addition of the fuselage generally moved $y_{c.p.}$ inboard approximately 3 percent of the semispan.\n\nPitching-Moment Characteristics\n\nNear zero lift the wing-alone aerodynamic center was located at 27 percent of the mean aerodynamic chord $\\left(\\frac{\\partial C_M}{\\partial C_L}\\right)_M = -0.02$ up to M = 0.80. This value compares with an estimated low-speed aerodynamic-center location of 24 percent $\\bar{c}$ (reference 3). The addition of the fuselage moved the aerodynamic center forward about 2 percent $\\bar{c}$ at the low Mach numbers.", "timestamp": "2026-07-22T06:17:28.491271+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 16, "total_pages": 47, "image_filename": "19930083221_p16.jpg", "text": "14\nNACA TN No. 1824\n\n$$\n\\overline{F}(s) = \\int_{0}^{\\infty} e^{-st} f(t) dt\n$$\n\nthen equation (21) can be rewritten in the form\n\n$$\n\\overline{C}_L(s) = s \\overline{C}_{L\\alpha}(s) \\overline{\\alpha}(s) \\tag{22}\n$$\n\nConsider now the case of a lifting flat plate oscillating harmonically without pitching at a frequency $\\omega$ and maximum angle of attack equal to $\\alpha_{max}$. Setting\n\n$$\n\\alpha(t) = \\alpha_{max} e^{i\\omega t} = \\alpha_{max} e^{i\\omega a_0 t'} \\tag{23}\n$$\n\nthen equation (22) yields\n\n$$\n\\frac{\\overline{C}_L(s)}{\\alpha_{max}} = \\frac{4}{s-i\\omega} \\left( \\sqrt{\\frac{2}{c_0 \\pi}} \\frac{e^{-s \\frac{c_0}{2}}}{\\sqrt{s}} - \\text{erf} \\sqrt{\\frac{c_0 s}{2}} \\right) \\tag{24}\n$$\n\nBy straightforward manipulation, the inverse transformation of equation (24) can be shown to give\n\n$$\n\\frac{C_L}{\\alpha_{max}} = 4 e^{i\\omega t} \\left\\{ 1 + \\sqrt{\\frac{2}{\\pi v}} e^{-iv} \\left[ C(\\omega t - v) - i S(\\omega t - v) \\right] \\right.\n$$\n\n$$\n\\left. - 2 \\sqrt{\\frac{2}{\\pi}} i N_1(\\omega t, v) - 2 \\sqrt{\\frac{2}{\\pi}} N_2(\\omega t, v) \\right\\} - \\frac{8}{\\pi} \\text{arc tan} \\sqrt{\\frac{2t-c_0}{c_0}} \\tag{25}\n$$\n\nwhere\n\n$$\nv = \\frac{\\omega c_0}{2} = \\frac{\\omega' c_0}{2a_0}\n$$", "timestamp": "2026-07-22T06:17:30.318604+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 15, "total_pages": 149, "image_filename": "19930083192_p15.jpg", "text": "NACA TN 1976\n\nThe instrument ranges were adjusted to suit the particular airplane. All instruments were equipped to give sufficient record time and were operated at film speeds to give adequate time resolution. A switch convenient to either the pilot or the observer enabled the instruments to be turned on or off at will.\n\nAll airplanes carried, in addition to the instruments mentioned, special instruments suited to the particular project. The XB-15 airplane, for example, was equipped with strain gages on wing beams, and optigraphs were provided to measure both wing and tail deflections. The XC-35 airplane carried four pressure recording units that were connected to orifices located on the wing of the airplane (see fig. 7) to measure the differential pressure between the upper and lower surfaces. Radiosonde data, as well as temperature measurements, were also obtained during the tests of the XC-35, XBM-1, Aeromca C-2, and F-61C airplanes. Table II gives the scope of the flight tests. With the exception of the flight investigations with the XB-15 airplane, all tests had objectives associated with research on gust structure, and the flight tests were under the control of the test engineer as to the performance and the type of flying done. The stability of the various airplanes was adjusted whenever possible so that the airplanes showed stick-fixed stability.\n\nRESULTS\n\nThe time histories of acceleration and airspeed obtained during these flight investigations were evaluated to obtain the effective gust velocity, the gradient distance, and the corresponding true velocity. As noted previously, every acceleration peak can be evaluated to obtain the effective gust velocity corresponding to the acceleration increment as measured from the 1 g datum, but only those acceleration peaks that are preceded by a smooth part of record can be evaluated to obtain the gust-gradient distance and the true gust velocity.\n\nGust intensity.- The available data on gust intensity, as measured by the effective gust velocities, have been examined in considerable detail in reference 9, and the frequency distributions given in figure 8 were derived therein. The results given in figure 8 are based on early German work, NACA flight tests, and V-G records from commercial transport operations. Curves A and B are the approximate limits of the frequency distributions of effective gust velocities for over-all operating conditions and show the probability that a gust will exceed any selected value of intensity.\n\nIn addition, a more detailed investigation of limited scope has been made (reference 10) of the XC-35 data to compare the frequency", "timestamp": "2026-07-22T06:17:31.507963+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 17, "total_pages": 62, "image_filename": "19930082918_p17.jpg", "text": "16\nNACA TN 1940\n\noccurred at the $\\frac{1}{2}$-hour age. Again the progressive improvement of\nunaged stock was evident with increasing rupture time. Figure 20 shows\nthat maximum deformation values (i.e., the elongation at the fracture)\nfor material aged at 1600° F followed the same pattern as for material\naged at 1400° F, with, however, shorter aging times required for an\nequivalent effect (compare curve for $\\frac{1}{2}$-hour aging, fig. 20, with\ncurve for 1-hour aging, fig. 16). Figure 21 shows that in no case\nwere the predominately intergranular failures, found in the specimens\naged a short time at 1400° F, observed in specimens aged at 1600° F\neven for as short a time as 0.5 hour.\n\nDISCUSSION OF RESULTS\n\nEffect of Aging on the Crystalline Structure\n\nUpon comparing the line-broadening, line peak-intensity, lattice-\nparameter, hardness, and metallographic data, it can be seen that:\n\n(1) Relatively low (111)-line intensity values resulting from the\naging of solution-treated low-carbon N-155 at 1400° F for time periods\nbetween approximately 0 and 30 hours were not accompanied by broadening\nof either the (111) or (220) line. It can thus be concluded that the\ndistortion causing the drop in peak intensity was of a short-period\nnature. When cognizance is taken of the fact that the lattice-parameter\ndata indicated that pronounced rejection of either interstitial or\nlarge-radius substitutional atoms did not take place until after\n10 hours or so of aging, it can be postulated that only nucleation\ntook place during the first 10 to 30 hours at 1400° F and that this\nnucleation process produced short-period strains. These time periods\nwill not, in general, be exact since each process tends to overlap\nthe next process to occur. The nuclei must be rather small since\nlarger nuclei would have a rather large spacing and any strains\nassociated would be long-period ones. Longer aging resulted in line-\nbroadening strains and, since this was associated with appreciable\nlattice contraction due to rejection of either interstitial or large-\nradius substitutional atoms, these strains were most probably associ-\nated with the actual precipitate particles. The metallographic data\nappear to bear these conclusions out since examination at magnifi-\ncations up to 10,000X revealed no evidence of precipitate particles\nuntil aging time periods at 1400° F were longer than 10 to 30 hours.\n\n(2) Relatively low (111)-line intensities during aging at 1600° F\nfor time periods up to 100 hours or so were associated with line\nbroadening. This indicated long-period strains. Further, the\nparameter measurements indicated an immediate lattice contraction by\nrejection of the precipitant atoms. Since, however, the metallographic", "timestamp": "2026-07-22T06:17:34.758791+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 90, "total_pages": 96, "image_filename": "19930085880_p90.jpg", "text": "88\nNACA RM No. L9C03\n\n18\nSpeed\n(fps) 30 25 20\n16\n14\n12\nTrimming moment, lb-ft\n10\n8\n6\n4\n2\n0\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n(d) $\\tau = 16^\\circ$.\nFigure 24.- Continued.\n15\n10\nNACA", "timestamp": "2026-07-22T06:17:34.966504+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 14, "total_pages": 37, "image_filename": "19930082646_p14.jpg", "text": "NACA TN 1980\n13\n\n[Figure: (a) Setup of model on towing apparatus.]\n\n(a) Setup of model on towing apparatus.\n\n[Figure: (b) Details of fore-and-aft gear. Labels: Rise slide wire, Fore-and-aft slide wire, Trim slide wire.]\n\n(b) Details of fore-and-aft gear.\n\nNACA\nL-59842\n\nFigure 3.- Model and towing apparatus.", "timestamp": "2026-07-22T06:17:40.642776+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 28, "total_pages": 65, "image_filename": "19930082546_p28.jpg", "text": "NACA TN No. 1870\n27\n\nREFERENCES\n\n1. Gütin, L.: Über das Schallfeld einer rotierenden Luftschraube.\nPhys. Zeitscher. der Sowjetunion, Bd. 9, Heft 1, 1936, pp. 57-71.\n\n2. Deming, Arthur F.: Propeller Rotation Noise Due to Torque and Thrust.\nNACA TN No. 747, 1940.\n\n3. London, Albert: Principles, Practice, and Progress of Noise Reduction\nin Airplanes. NACA TN No. 748, 1940.\n\n4. Nichols, R. H., Jr., Sleeper, H. P., Jr., Wallace, R. L., Jr., and\nEricson, H. L.: Acoustical Materials and Acoustical Treatments\nfor Aircraft. Jour. Acous. Soc. Am., vol. 19, no. 3, May 1947.\n\n5. Davis, A. H.: Modern Acoustics. G. Bell and Sons Ltd. (London),\n1934.\n\n6. Den Hartog, J. P.: Mechanical Vibrations. Second ed., McGraw-Hill\nBook Co., Inc., 1940.", "timestamp": "2026-07-22T06:17:43.958061+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 11, "total_pages": 32, "image_filename": "19930085847_p11.jpg", "text": "NACA RM A9D04\n\nCONFIDENTIAL\n\nTEST\n\n485299\n\nCONFIDENTIAL\n\nNACA\nA-12653\n\nFigure 1.- Test airplane as instrumented for the flight tests.\n\n9", "timestamp": "2026-07-22T06:17:44.604762+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 49, "total_pages": 114, "image_filename": "19930086061_p49.jpg", "text": "```markdown\nNACA RM L59D07\n\n| | $\\alpha = 34.1^\\circ$ <br> $C_L = 1.12$ | $\\alpha = 39.1^\\circ$ <br> $C_L = 1.17$ | $\\alpha = 44.1^\\circ$ <br> $C_L = 0.94$ | $\\alpha = 50.1^\\circ$ <br> $C_L = 0.62$ |\n| :--- | :---: | :---: | :---: | :---: |\n| **Station 1** <br> $\\frac{y}{b/2} = 0$ | [Graph: P vs x/c] <br> -4 to 1 (P) <br> 0 to 10 (x/c) | [Graph: P vs x/c] <br> Station 1 <br> $\\frac{y}{b/2} = 0$ | [Graph: P vs x/c] <br> [Diamond shape] | [Graph: P vs x/c] |\n| **Station 2** <br> $\\frac{y}{b/2} = 0.167$ | [Graph: P vs x/c] <br> -4 to 1 (P) <br> 0 to 10 (x/c) <br> Legend: <br> - - - Upper <br> - - - Lower | [Graph: P vs x/c] <br> Station 2 <br> $\\frac{y}{b/2} = 0.167$ | [Graph: P vs x/c] | [Graph: P vs x/c] |\n| **Station 3** <br> $\\frac{y}{b/2} = 0.333$ | [Graph: P vs x/c] <br> -3 to 1 (P) <br> 0 to 10 (x/c) | [Graph: P vs x/c] <br> Station 3 <br> $\\frac{y}{b/2} = 0.333$ | [Graph: P vs x/c] | [Graph: P vs x/c] <br> NACA |\n\n(a) Stations: 1, 2, 3.\n\nFigure 11.- Chordwise pressure distribution about wing 3 at angles of attack of $34.1^\\circ$, $39.1^\\circ$, $44.1^\\circ$, and $50.1^\\circ$; $\\psi = 0^\\circ$.\n\n45\n```", "timestamp": "2026-07-22T06:17:47.472792+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 23, "total_pages": 28, "image_filename": "19930082703_p23.jpg", "text": "NACA TN 1983\n\n21\n\n[Figure: A black-and-white photograph of a helicopter on a paved surface, viewed from the side. The helicopter has a star insignia on its fuselage and a long main rotor blade extending to the left. A small label with \"NACA\" and \"I-59515\" is visible on the right side of the image.]\n\nFigure 4.- Helicopter C.", "timestamp": "2026-07-22T06:17:49.783776+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 29, "total_pages": 46, "image_filename": "19930082613_p29.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:17:51.662748+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 16, "total_pages": 17, "image_filename": "19930085572_p16.jpg", "text": "14\nNACA RM No. E8L02\n\nPressure\naltitude\n(ft)\no 5,000\n□ 10,000\n◇ 20,000\n△ 30,000\n\nFuel\n—○— AN-F-58\n- -●- - AN-F-32\n\nCorrected tail-pipe temperature, °R\n2200\n2000\n1800\n1600\n1400\n1200\n1000\n\n3000 4000 5000 6000 7000 8000 9000\nCorrected engine speed, rpm\n\n[NACA logo]\n\nFigure 5. - Comparison of corrected tail-pipe temperature with AN-F-58\nand AN-F-32 fuels. Mach number, 0.37.", "timestamp": "2026-07-22T06:17:56.140794+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 32, "total_pages": 44, "image_filename": "19930082566_p32.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:17:57.599122+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 16, "total_pages": 72, "image_filename": "19930085491_p16.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 15\n\n| Configuration | Reynolds No. x $10^{-6}$ | Figure No. |\n| :--- | :--- | :--- |\n| WF-57 | 0.62 | 7(a) |\n| WF-60 | 0.62 | 7(b) |\n| WF-63 | 0.31, 0.62, 0.84 | 7(c) |\n| WF-67 | 0.62, 0.95 | 7(d) |\n| WF-70 | 0.62 | 7(e) |\n| Fuselage alone$^1$ | 0.62 | 7(f) |\n\nAlso shown on these figures are the theoretical characteristics wherever they were determined.\n\nFigure 8 shows to a larger scale the pitching-moment data and also includes the position of the center of pressure plotted against lift coefficient. Figure 9 is a replot of the moment data in which the moments are referred to the quarter chord, rather than the mid-chord of the mean aerodynamic chord. These data may be used directly in comparing the pitching-moment characteristics with those obtained in the subsonic investigations.\n\nFigure 10 presents cross plots of the major aerodynamic and geometric parameters against the factor m, which is the ratio of the cotangent of the sweep angle of the leading edge to the cotangent of the sweep angle of the Mach line. The variations of minimum drag coefficient, lift-curve slope, drag-rise factor, and maximum lift-drag ratio are shown in figures 10(a) through 10(d), respectively, for a Reynolds number of 0.62 million. Figure 11 presents the variations of $\\Delta C_D/(\\Delta C_L)^2$, $k_a$, and $\\Delta C_L/\\Delta \\alpha$ with lift coefficient for WF-63 at a Reynolds number of 0.62 million.\n\nTable II summarizes the results of the force tests for all configurations and Reynolds numbers investigated. In cases where theoretical values have been calculated they have been entered in parentheses directly below the experimental value. The theoretical results, based on linear theory give straight-line lift and moment curves and parabolic drag curves. The experimental results, however, in all cases show nonlinear lift and moment curves and drag curves which are composed essentially of two parabolic segments that intersect at slightly less than the optimum lift coefficient. Because of these variations, two values of $dC_L/d\\alpha$, $\\Delta C_D/(\\Delta C_L)^2$, and $k_a$ are shown in table II for each configuration, the values being those for zero lift and for the optimum lift coefficient. The variation of these parameters with lift coefficient will be considered subsequently.\n\n$^1$Reynolds number and coefficients are based on reference lengths and area of WF-63.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:17:59.343048+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 5, "total_pages": 60, "image_filename": "19930085862_p5.jpg", "text": "NACA RM No. L9A07\n\n$C_{N_a}$ aileron normal-force coefficient (Aileron normal force/$qS_a$)\n\n$C_{h_a}$ aileron hinge-moment coefficient  \n(Aileron hinge moment about hinge line/$qb_a\\bar{c}_a^2$)\n\n$P_R$ aileron balance-chamber resultant-pressure coefficient  \n((Lower-surface pressure - Upper-surface pressure)/q)\n\n$q$ free-stream dynamic pressure, pounds/square foot\n\n$b$ wing span measured normal to plane of symmetry, feet\n\n$b_s$ spoiler span measured normal to plane of symmetry, feet\n\n$b_a\\bar{c}_a^2$ product of aileron span, measured along aileron hinge line,  \nand square of root-mean-square chord, measured behind and  \nnormal to hinge line, 0.536 cubic feet\n\n$S$ wing area, square feet\n\n$S_a$ aileron area behind hinge line, square feet\n\n$\\bar{c}$ wing mean aerodynamic chord measured parallel to plane of  \nsymmetry, 2.942 feet $\\left(\\frac{2}{S}\\int_0^{b/2} c^2 dy\\right)$\n\n$c$ local wing chord measured parallel to plane of symmetry, feet\n\n$\\bar{c}_b$ root-mean-square chord of hypothetical aileron balance measured  \nahead of and normal to aileron hinge line, feet\n\n$y$ spanwise coordinate, measured normal to plane of symmetry, feet\n\n$\\alpha$ angle of attack, degrees\n\n$\\delta_a$ aileron deflection, measured in plane normal to hinge line,  \ndegrees (positive when trailing edge is deflected downward)\n\n$\\delta_{a_{total}}$ arithmetical sum of equal up and down aileron deflections for  \nan assumed set of ailerons\n\n$\\Lambda$ sweepback of leading edge of wing, degrees", "timestamp": "2026-07-22T06:18:00.057292+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 14, "total_pages": 46, "image_filename": "19930085548_p14.jpg", "text": "NACA RM No. E8L30 13\n\nwhere\n\n$H_g$ enthalpy of exhaust gas, Btu per pound\n\n$Q_t$ heat rejection, fraction of heat input neglecting friction\n\n$\\eta_t$ indicated thermal efficiency from experimental data\n\n$h_c$ heat of combustion of fuel, 18,500 Btu per pound\n\nF/A fuel-air ratio\n\n$H_m$ enthalpy of inlet air, Btu per pound\n\nData from reference 9 allowed graphical expression of these enthalpy values as functions of temperature and fuel-air ratio and thus permitted a solution of the equation to be made. Equation (5) results in the same exhaust-gas temperature as that calculated in the analysis in reference 1 for a corresponding heat-rejection rate.\n\nValues of exhaust-gas temperature calculated by means of equation (5) are compared with the experimentally determined values in figure 21. The calculated values appear to be in good agreement with the experimental results. At a fuel-air ratio of 0.04 and beyond, all the data points lie below the calculated curves. This disparity may be a result of decreasing combustion efficiency in the experimental data because the equations assume 100-percent combustion efficiency, or of neglected heat losses at the high temperatures involved.\n\nThe variation in calculated and experimental exhaust-gas temperatures with changes in compression ratio is shown in figure 22. Here again the correlation is shown to be good. The change in exhaust-gas temperature with compression ratio is caused by variation in the thermal efficiency of the cylinder.\n\nEFFECT OF DIFFERENCES BETWEEN OBSERVED AND PREDICTED PERFORMANCE\n\nInasmuch as this investigation was conducted to determine experimentally the performance characteristics of a cylinder for gas-generator use and to compare the data obtained with those assumed for the previous analysis of the gas generator (reference 1), examination of the effect of differences in the experimental and assumed data on the performance of the gas-generator engine is of interest.", "timestamp": "2026-07-22T06:18:02.721707+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 31, "total_pages": 50, "image_filename": "19930082592_p31.jpg", "text": "**Page intentionally left blank**\n\n**Page intentionally left blank**", "timestamp": "2026-07-22T06:18:02.729514+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 17, "total_pages": 47, "image_filename": "19930083221_p17.jpg", "text": "NACA TN No. 1824\n15\n\n$\\omega'$ is true impressed frequency\n($\\alpha = \\alpha_{max} e^{i\\omega't'}$).\n\n$C(\\omega t - \\nu)$, $S(\\omega t - \\nu)$ are Fresnel's\nintegrals (reference 11).\n\n$N_1(\\omega t, \\nu)$, $N_2(\\omega t, \\nu)$ are integrals\ndefined in the appendix.\n\nIf the response to a cosine\nvariation of $\\alpha$ is required,\nonly the real part of equation\n(25) is used. Such a response\nin the early stages of the\nmaneuver is shown in figure 5.\nFor very large values of time\n$C_L$ as given by equation (25)\napproaches the value\n\n[Figure: Graph showing $C_L/\\alpha$ vs $\\omega t$. Curves labeled $C_L$ and $\\alpha$. X-axis: 0, 2, 4, 6, 8, 10 $\\omega t$. Y-axis: -4, -2, 0, 2, 4.]\n\nFigure 5.— Lift resulting from\ncosine-wave angle-of-attack\nvariation.\n\n$$ \\frac{C_L}{\\alpha_{max}} = 4e^{-i\\omega t} \\left( \\sqrt{\\frac{1}{i\\pi\\nu}} e^{i\\nu} - \\text{erf} \\sqrt{i\\nu} \\right) \\quad (26) $$\n\nfrom which both the amplitude and phase shift of $C_L$ resulting from\neither an impressed sine or cosine variation of $\\alpha$ can be readily\n\n[Figure: Graph showing $(C_L)_{MAX}/(\\alpha)_{MAX}$ vs $\\omega c_0/2$. Y-axis: 0, 3, 4, 5, 6. X-axis: 1, 2, 3, 4, 5, 6, 7, 8, 9. Annotations: \"initial value\" pointing to the start of the curve, \"final value\" pointing to the asymptotic part of the curve.]\n\nFigure 6.— Amplitude of oscillatory lift resulting from a cosine\nangle-of-attack oscillation (without pitching) at $M_0 = 1$.", "timestamp": "2026-07-22T06:18:04.931243+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 15, "total_pages": 37, "image_filename": "19930082646_p15.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:18:07.193386+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 91, "total_pages": 96, "image_filename": "19930085880_p91.jpg", "text": "NACA RM No. L9C03\n89\n\n[Figure: A graph plotting Trimming moment against Wetted area. The graph contains multiple data series represented by different symbols (circles, squares, diamonds, triangles, inverted triangles) corresponding to different speeds. A legend box indicates the symbol for Speed (fps).]\n\nTrimming moment, lb-ft\n18\n16\n14\n12\n10\n8\n6\n4\n2\n0\n\nSpeed\n(fps)\n30\n25\n20\n15\n10\n\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n(e) $\\tau = 20^\\circ$.\nFigure 24.- Concluded.\nNACA", "timestamp": "2026-07-22T06:18:10.229308+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 9, "total_pages": 31, "image_filename": "19930085859_p9.jpg", "text": "NACA RM No. L9B25\n\nAt $C_L = 0.4$ the wing-alone aerodynamic center was about 25 percent $\\bar{c}$ at low Mach numbers and moved back to 46 percent $\\bar{c}$ at the highest Mach numbers. The destabilizing effect of the fuselage was slightly more pronounced at $C_L = 0.4$ than at $C_L = 0$.\n\nDownwash and Dynamic-Pressure Surveys\n\nThe variation of effective downwash angle with tail height and angle of attack for the wing alone and wing-fuselage at various Mach numbers is presented in figures 9 and 10. The downwash gradient $\\partial \\epsilon / \\partial \\alpha$ near zero lift for the wing alone (fig. 11) increased as the tail location approached the chord plane, at Mach numbers below 1.00. Above $M = 1.00$ $\\partial \\epsilon / \\partial \\alpha$ was maximum at a tail location of 30 percent semispan below the chord plane. At the higher lift coefficients $\\partial \\epsilon / \\partial \\alpha$ was generally less than the zero lift value for tail positions below the chord plane and was higher for tail positions above the chord plane.\n\nThe addition of the fuselage caused a marked increase in $\\partial \\epsilon / \\partial \\alpha$ for tail positions near the chord plane (figs. 10 and 11) up to $M = 0.95$. Above $M = 1.00$ the effect of the fuselage on the downwash gradient near the chord plane was small. Note that the test angle-of-attack range with the free-floating tails nearest the chord line extended was restricted because of the presence of the fuselage.\n\nThe results of point dynamic-pressure surveys made in a vertical plane containing the 25-percent mean-aerodynamic-chord point of the free-floating tails used in the downwash surveys are presented in figure 12. The maximum loss in dynamic pressure at the wake center line for the higher angles of attack was never more than 15 percent of the free-stream dynamic pressure.\n\nThe addition of the fuselage showed practically no effect on the dynamic-pressure ratios throughout most of the Mach number range. At $10^\\circ$ angle of attack at the higher Mach numbers the addition of the fuselage shifted the wake center line above that of the wing alone.", "timestamp": "2026-07-22T06:18:10.578756+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 18, "total_pages": 62, "image_filename": "19930082918_p18.jpg", "text": "NACA TN 1940\n\ndata did not indicate that visible precipitate particles appeared much before 100 hours at $1600^\\circ$ F had elapsed, these long-period strains were associated with large precipitant nuclei with large spacing. The spacing of the stable nuclei was certainly larger when aging was carried on at $1600^\\circ$ F than when aging was done at $1400^\\circ$ F as evidenced by the relative spacing of the precipitate particles which finally appeared at the two temperatures. That the stable nuclei were probably larger at $1600^\\circ$ F than at $1400^\\circ$ F is in agreement with Mehl and Jetter's summaries in regard to precipitation from solid solution (see reference 7).\n\n(3) The data for aging at $1200^\\circ$ F were only fragmentary but in view of the very long period (approx. 1000 hr) during which no appreciable lattice contraction occurred and very little visible precipitate appeared, one can conclude that only short-period nucleation occurred during the aging time studied. The changes in (111)-line intensity were then associated with the short-period strains of nucleation. The two separate minimums found could possibly be associated with, first, the matrix material in contact with the boundary and, second, the interior matrix nucleation.\n\nThe hardness data appeared to correlate quite generally with line broadening — the greater the degree of line broadening, the greater the hardness. Hardness, then, as far as the alloy studied is concerned, was associated with long-period strains. In other words the long-period strains, associated with the precipitate particles formed after the nucleation period, increased deformation resistance under the conditions of large localized deformation present during a hardness test.\n\nFactors Controlling Creep Strength\n\nComparison of the creep rates at 30,000 psi for materials aged at either $1200^\\circ$, $1400^\\circ$, or $1600^\\circ$ F (see fig. 14) with the results of the structure measurements showed that the loss of initial creep strength at this stress level was most clearly associated with matrix depletion of the relatively large-radius or the interstitial precipitant atoms. This was also associated with:\n\n(1) Removal of the short-period nucleation strains in the case of material aged at $1400^\\circ$ F.\n\n(2) Development of visible precipitate and the continuous, relatively wide, grain boundary phase. These precipitant particles and the grain boundary phase were initially surrounded by concentration gradients which became progressively less steep with increased aging time (and/or increased aging temperature). (See figs. 5 to 9.)", "timestamp": "2026-07-22T06:18:10.734944+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 16, "total_pages": 149, "image_filename": "19930083192_p16.jpg", "text": "12\nNACA TN 1976\n\ndistributions of effective gust velocity in thunderstorms and line squalls for various altitude ranges. Table III summarizes the data obtained to a threshold of 4 feet per second for several ranges of altitude from the surface to 35,000 feet. Table III also lists the record time and the average spacing between gusts for each altitude range. Figure 9 shows the probability distributions of effective gust velocity for the several altitude ranges that are obtained by curves fitted to the data of table III. In this investigation, the data for altitudes below 5000 feet were not used because the conditions investigated represented random record taking in both clear air and clouds at the end of the flights.\n\nFrom the tests performed with the XC-35 and the F-61C airplanes, the variation of expected maximum effective gust velocity with altitude within convective clouds has been determined. These results are summarized in figure 10, in which the data for the XC-35 airplane correspond to 100 miles of rough-air flying, while those for the F-61C airplane correspond to 1000 miles. The latter data were obtained during the 1946 operations of the U.S. Weather Bureau thunderstorm project at Orlando, Fla., and are based on the maximum values of effective gust velocity for each 3000-foot interval; however, the XC-35 data utilized a complete count of all gusts.\n\nGust spacing.- Some information on gust spacing, that is, the distance from peak to peak, has also been obtained. In reference 11, an analysis was made of the XC-35 data to determine the average spacing between the large gusts (defined as $U_g$ greater than 5 to 8 fps) in areas of continuous rough air. Continuous rough air is defined as a sequence of large gusts in which the spacing between gusts was less than 2.2 seconds. Table IV, reproduced from reference 11, summarizes some of the pertinent characteristics of the large isolated and repeated gusts. The table presents a comparison of the average and the range of intensities and the average and range of spacings for isolated single gusts and sets of two and three repeated gusts occurring with equal frequency.\n\nMore general information on gust spacing, contained in reference 9, indicates that the average spacing between gusts that were counted to an estimated threshold of 0.3 foot per second was a function of the airplane mean geometric chord. This count indicated an average spacing between successive peaks of 11 chords so that the number of gusts per mile of rough air is roughly equal to 500/$\\bar{c}$. The definition of the number of gusts per mile of rough air led to the concept of \"path ratio\" which provides a measure of the percent of rough air encountered under operating conditions.\n\nGust-gradient distance.- Figure 11 presents a typical plot of measurements of gradient distance and gust velocity. Inspection of the", "timestamp": "2026-07-22T06:18:15.592888+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 24, "total_pages": 28, "image_filename": "19930082703_p24.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:18:20.697724+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 17, "total_pages": 17, "image_filename": "19930085572_p17.jpg", "text": "NACA RM No. E8L02\n15\n\nFuel\n—□— AN-F-58\n--○-- AN-F-32\n\nPressure altitude, ft\n30,000\n25,000\n20,000\n15,000\n10,000\n5,000\n0\n\nActual engine speed, rpm\n0\n1000\n2000\n3000\n4000\n\n[Figure: Graph showing pressure altitude vs. actual engine speed for two fuel types, with data points and trend lines. NACA logo in bottom right corner of graph area.]\n\nFigure 6. - Comparison of effect of altitude on combustor blow-out limits with AN-F-58 and AN-F-32 fuels. Mach number, 0.37.", "timestamp": "2026-07-22T06:18:21.910082+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 33, "total_pages": 44, "image_filename": "19930082566_p33.jpg", "text": "NACA TN NO. 1889\n\nMaximum principal stress, $\\sigma_1'$, psi\n\nN, cycles\n\n(a) For stress ratio $R = \\sigma_2' / \\sigma_1' = 0$.\n\nFigure 12.- S-N curves.\n\n31", "timestamp": "2026-07-22T06:18:24.489927+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 29, "total_pages": 65, "image_filename": "19930082546_p29.jpg", "text": "28\nNACA TN No. 1870\n\n[Figure: Diagram showing a sphere with coordinate axes $x, y, z$ and $x', y', z'$. Angles $\\theta, \\chi, \\gamma, \\delta$ are indicated. A line labeled \"Observer\" passes through point B. The NACA logo is present.]\n\nFigure 1.- Description of coordinate system.", "timestamp": "2026-07-22T06:18:31.520023+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 30, "total_pages": 46, "image_filename": "19930082613_p30.jpg", "text": "NACA TN 1938\n29\n\n[Figure: Micrograph of a cracked metal specimen with labels pointing to various features. A stamp in the upper right corner reads: NACA C-22628 12-9-48]\n\nSubsurface scale\n(Internal oxide)\n\nCrack\n\nMain crack\n\nSolid-surface\nscale\n\nMetal\n\nThis branch is probably\na tubular formation of\ninternal oxide around\ncrack. (fig. 13)\n\n(b) Detail A; magnification, X1500.\n\nFigure 8. - Concluded. Portion of large crack showing surface and subsurface scale. Type-A liner; etchant, none; condition, failed in service.", "timestamp": "2026-07-22T06:18:35.548838+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 12, "total_pages": 32, "image_filename": "19930085847_p12.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:18:36.008304+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 32, "total_pages": 50, "image_filename": "19930082592_p32.jpg", "text": "NACA TN 1914\n31\n\n[Figure: Micrograph showing a circular cross-section of a material. The top half is dark and labeled \"Oxide\". The bottom half is lighter and speckled, labeled \"Unoxidized ceramal\". A line between them is labeled \"Oxidation interface\". In the bottom right corner of the figure is a NACA logo with the text \"C-22909\" and \"2-7-49\".]\n\nFigure 10. - Oxidation interface of 30-percent-molybdenum - titanium carbide ceramal. Oxidation proceeds as linear front without grain-boundary penetration. Temperature, 1785° F; time at temperature, 7 hours; etchant, potassium hydroxide plus potassium ferricyanide KOH+K$_3$Fe(CN)$_6$; magnification, X500.", "timestamp": "2026-07-22T06:18:38.663162+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 50, "total_pages": 98, "image_filename": "19930086073_p50.jpg", "text": "48\n\nLift coefficient, $C_L$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:18:39.803665+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 17, "total_pages": 72, "image_filename": "19930085491_p17.jpg", "text": "16 CONFIDENTIAL NACA RM No. A8J04\n\nLiquid-Film Tests\n\nThe results of the liquid-film tests are presented in the form of line drawings, since inconsistent lighting effects throughout the investigation resulted in a nonuniform set of test photographs. Two photographs are included, however, which show typical results obtained at zero lift and near the optimum lift coefficient. These are shown in figure 12 with the corresponding line drawings.\n\nAt zero lift (fig. 12(a)), the pattern obtained on the wing of WF-63 revealed that the boundary-layer flow on the inboard sections was laminar back to the trailing edge with the exception of a small turbulent flow area close to the body that originated at the juncture of the fuselage and wing leading edge. On the outboard sections, the boundary layer was laminar back to approximately 60 percent of the chord where the flow separated from the wing surface. The separation line is indicated in the photograph by a ridge of fluid on the surface which results from the opposing shear forces acting on the liquid film ahead of and behind the line of separation.\n\nThe lifting wing upper-surface photograph shown in figure 12(b) reveals that the laminar boundary layer separates closer to the wing leading edge than at zero lift. After separating, however, the boundary layer reattaches as a turbulent boundary layer on the inboard sections as is evidenced by the drying lines behind the separated region. The outward curvature of these lines indicated a spanwise boundary-layer flow. Outboard of the section where the line of reattached flow intersects the trailing edge of the wing, the photograph shows evidence of a secondary flow within the stalled region. On a later test run with only the bottom surface of the wing coated with liquid film, this pattern was observed to result from air flow around the trailing edge into the upper-surface separated region. In the absence of pressure-distribution studies in these tests the reason for the formation of a fluid ridge within this separated region is not immediately apparent. However, like the line of laminar separation, it must occur where there is zero surface shear in the chordwise direction.\n\nAll boundary-layer-flow drawings are of the upper wing surfaces. Since the wings were symmetrical, the patterns obtained on the bottom surfaces at zero lift were the same as those for the upper surfaces. Where upper-surface patterns are presented for the lifting wings, the lower surfaces were observed to have completely laminar boundary-layer flow.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:18:41.207574+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 18, "total_pages": 47, "image_filename": "19930083221_p18.jpg", "text": "16\nNACA TN No. 1824\n\ndetermined. Figure 6 shows the amplitude of lift oscillation corre-\nsponding to continuous angle-of-attack oscillation, as determined\nfrom equation (26), plotted as\na function of $v = \\frac{\\omega c_0}{2}$. It is\napparent that as $\\omega$ approaches\nzero, $C_L$ approaches infinity;\nthat is, as the impressed wave\napproaches the \"step\" function\n(fig. 7), the lift coefficient\napproaches infinity. This\nresult is in agreement with\nequation (20b) as t approaches\ninfinity. As the frequency\nparameter V is increased,\nhowever, the value of $C_{L_{max}}/\\alpha_{max}$\n\n<!-- Image (112, 152, 474, 319) -->\n\nFigure 7.- Variation of $\\cos \\omega t$\nwith t for various values of\n$\\omega$.\n\nis reduced and reaches a minimum of about 3.4 for a value of $v = 0.9$.\nFor a speed of sound around 1000 feet per second and a wing chord of\n6 feet, this would correspond to a frequency of 47.7 cycles per second,\na value well within the range of practical flutter frequencies. It is\ninteresting to note that figure 6 also shows that as the frequency of\noscillation becomes large (i.e., $v>3$) the value of $C_{L_{max}}/\\alpha_{max}$\napproaches the value 4. This is the same as the value for $C_{L\\alpha}(t)$\nin the early stages following a step variation of $\\alpha$.\n\nUnsteady State, $M_0 < 1$\n\nIt has been pointed out that, in the determination of the\nindicial lift function for a wing traveling at subsonic speeds, the\nlifting-surface analogue involves the calculation of load distribution\nover a swept-forward wing with subsonic edges. This means that\nrecourse cannot be made in the solution to the simple source distri-\nbution method used in reference 7 to treat the $M_0 \\ge 1$ case. Since,\nhowever, a portion of the leading edge in the present case is still\nsupersonic, the problem is particularly adapted to lifting-surface\nmethods developed by Evvard in reference 12. Figure 8 indicates, as\nin figure 1(b), the geometry associated with the boundary conditions.\nIn the Evvard analysis the solutions, as in the previous case, are\ncalculated for various regions. As an aid in identifying the different\nresults the sketch denotes these regions by Roman numerals.", "timestamp": "2026-07-22T06:18:49.123098+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 10, "total_pages": 31, "image_filename": "19930085859_p10.jpg", "text": "8\nNACA RM No. L9B25\n\nThe dynamic-pressure surveys show that for the particular tail\nlength used a tail position of 10 percent of the semispan or more below\nthe chord plane would generally be most favorably located from consider-\nation of wake effects.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nREFERENCES\n\n1. Weil, Joseph, and Goodson, Kenneth W.: Aerodynamic Characteristics\nof a Wing with Quarter-Chord Line Swept Back 45°, Aspect Ratio 4,\nTaper Ratio 0.6, and NACA 65A006 Airfoil Section. Transonic-Bump\nMethod. NACA RM No. L9A21, 1949.\n\n2. Schneiter, Leslie E., and Ziff, Howard L.: Preliminary Investigation\nof Spoiler Lateral Control on a 42° Sweptback Wing at Transonic\nSpeeds. NACA RM No. L7F19, 1947.\n\n3. DeYoung, John: Theoretical Additional Span Loading Characteristics\nof Wings with Arbitrary Sweep, Aspect Ratio, and Taper Ratio.\nNACA TN No. 1491, 1947.", "timestamp": "2026-07-22T06:18:51.076784+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 92, "total_pages": 96, "image_filename": "19930085880_p92.jpg", "text": "90\nNACA RM No. L9C03\n\nSpeed,\n(fps)\nO 10\n□ 15\n◇ 20\n△ 25\n▽ 30\n\nDraft, ft\n.64\n.56\n.48\n.40\n.32\n.24\n.16\n.08\n0\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n\n(a) $\\tau = 4^0$.\n\nFigure 25.- Variation of draft with wetted area. Model 250D.", "timestamp": "2026-07-22T06:18:53.349665+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 25, "total_pages": 28, "image_filename": "19930082703_p25.jpg", "text": "NACA TN 1983\n23\n\nStick\nmotion\nfrom trim,\nin.\n\nRearward 5\n0\nForward 5\n\n- Rearward stop\n- Forward stop\n\nPitching\nvelocity,\ndeg/sec\n\nNose up 20\n0\nNose down 20\n\nNormal\nacceleration,\ng\n\n1.5\n1.0\n.5\n0\n\n0 2 4 6 8\nTime, sec\n\n[NACA logo]\n\nFigure 5.- Time history of a pull-up maneuver for helicopter A\nat 80 miles per hour.\n\nStick position,\ninches forward\n\n0\n5\n\nNormal\nacceleration,\ng\n\n1.5\n1.0\n.5\n0\n\n0 5 10 15 20\nTime, sec\n\n- Forward stop\n\n[NACA logo]\n\nFigure 6.- Time history of an attempted helicopter oscillation which\nrequired a pull-up for recovery. (From reference 1.)", "timestamp": "2026-07-22T06:18:59.678982+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 16, "total_pages": 37, "image_filename": "19930082646_p16.jpg", "text": "NACA TN 1980\n15\n\nWarped forebody and extended afterbody ———\nBasic forebody and basic afterbody - - - - -\n\n[Figure: Graph plotting Trim, deg (y-axis, 0-12) against Speed, mph (x-axis, 40-90). The graph contains data points represented by triangles and lines indicating stability limits. Regions are labeled \"Unstable\" and \"Stable\". A NACA logo is present in the bottom right corner of the plot area.]\n\nFigure 4.- Trim limits of stability.", "timestamp": "2026-07-22T06:18:59.857603+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 17, "total_pages": 149, "image_filename": "19930083192_p17.jpg", "text": "NACA TN 1976\n13\n\nfigure indicates that the scattered pattern of data is not amenable to detailed analysis and, for such purposes, the data on gust-gradient distance were sorted according to the variable under consideration. Figures 12(a), 12(b), and 12(c) present the data on the average gust-gradient distance as a function of the gust velocity (width of bracket, 4 fps). Figure 12(a) is a summary of the data for all airplanes with $H_{av}$ expressed as multiples of the mean geometric chord; figure 12(b) presents the same data with $H_{av}$ expressed in multiples of the airplane wing span; and figure 12(c), in feet. For comparison with other figures and for subsequent use, a faired curve was drawn through the data in figure 12(a) to agree best with the data from the XC-35 airplane.\n\nThe results presented in figures 12(a), 12(b), and 12(c) do not include the complete data for every airplane since some brackets, particularly below about 8 feet per second, are not complete. The magnitude of the acceleration increments for values of U less than 8 feet per second was about the same order as the resolution of the accelerometer. For the larger gust velocities, some \"average\" values were not plotted since the number of points represented was less than three.\n\nIn some cases, sufficient data were available to obtain the probable values of the gust-gradient distance. The probable value was determined by fitting a theoretical frequency distribution to the data and determining the point with the highest frequency. The results are plotted against the corresponding average gust-gradient distance $H_{av}$ in figure 13. Figure 13 includes two solid lines, a 45° line representing the equality of the quantities and a line through the data offset by 2 chords.\n\nData for the gust-gradient distance for the XC-35 and F-61C airplanes have been utilized to determine the variations of the average gradient distance expressed in chords as a function of altitude. (See fig. 14.) In the case of the XC-35 airplane where no fixed altitudes were flown, altitude brackets were selected, and the average gradient distance was plotted against the center value for the pertinent bracket. The F-61C airplane tests were made at fixed altitudes, and the data have been plotted at the appropriate altitude level.\n\nSpanwise gust distribution.- Records of local wing pressures and airspeeds obtained during the flight investigation with the XC-35 airplane (reference 12) were evaluated to obtain the local values of effective gust velocity at each of four spanwise stations. The general method followed was to plot the data for each gust as a function of the spanwise location and connect the points by straight lines. The individual spanwise distributions were later identified with the corresponding values of the gust velocity U and the effective gust velocity $U_e$ obtained from the other flight records.", "timestamp": "2026-07-22T06:19:00.341255+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 30, "total_pages": 65, "image_filename": "19930082546_p30.jpg", "text": "NACA TN No. 1870\n29\n\n[Figure: Four propeller test blades of varying shapes and sizes, arranged side-by-side on a flat surface against a dark background.]\n\nNACA\nL-56022\n\nFigure 2.- Propeller test blades.", "timestamp": "2026-07-22T06:19:02.437980+00:00"}